1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a transition duct positioned between the combustor and the turbine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, includes a combustor that produces a hot gas flow, a multiple stage turbine that extracts mechanical energy from the hot gas flow by producing rotation of the rotor shaft, and a transition duct positioned between the combustor and the turbine to direct the hot gas flow into the turbine section. The combustor section could be a single annular combustor or a plurality of can combustors arranged annularly around the engine.
In the multiple can combustor arrangement, each can combustor is associated with a transition duct. The prior art U.S. Pat. No. 6,890,148 B2 issued to Nordlund on May 10, 2005 and entitled TRANSITION DUCT COOLING SYSTEM shows one of these transition ducts with a circular inlet on the combustor end and a rectangular outlet with an arched configuration on the outlet. A plurality of these transition ducts are arranged around the engine to form an annular outlet leading into the turbine section. In this type of engine, a separate stator vane assembly is secured to the engine between the transition ducts and the turbine inlet.
Several prior art references include a guide vane assembly within the transition duct to avoid the expensive separate production and assembly in addition to the subassemblies of each combustion chamber. U.S. Pat. No. 5,953,919 issued to Meylan on Sep. 21, 1999 and entitled COMBUSTION CHAMBER HAVING INTEGRATED GUIDE BLADES discloses a transition duct with guide blades built into the duct at the end. Other patents that show guide vanes formed with the transition duct are U.S. Pat. No. 2,630,679 issued to Sedille on Mar. 10, 1953 and entitled COMBUSTION CHAMBER FOR GAS TURBINES WITH DIVERSE COMBUSTION AND DILUENT AIR PATHS; and U.S. Pat. No. 3,316,714 issued to Smith et al on May 2, 1967 and entitled GAS TURBINE ENGINE COMBUSTION EQUIPMENT.
One major problem with the above identified prior art transition ducts is that the guide vanes, which are exposed to the highest gas flow temperature within the engine, are thermally coupled to the duct, and as a result experience very high thermal gradients that lead to very high stress levels. This shortens the life of the guide vanes and the portions of the duct that secure the guide vanes. Also, the transition ducts of the prior art do not allow for the capability of airfoils that are made from a single crystal material as in the present invention.